This invention relates generally to the field of turbines and power systems. Fuel fired internal combustion engines such as gas turbine engines utilize a working fluid, namely an air/fuel mixture, which changes composition during combustion to drive the turbine with hot expanded gases. A conventional gas turbine engine includes a compressor, a combustion chamber and a turbine made up of an arrangement of stators and rotors. Each of the rotors includes blades and a supporting disk. Ideally, for the optimum extraction of energy, the combustion process should occur at about 4000xc2x0 Fahrenheit. However, as a practical matter due to metallurgical concerns, the components of a turbine must operate at considerably lower temperatures. Cooling of the stationary housing and stators in combustion chamber walls is relatively straightforward by any of a number of means; however the rotors, due to their high rotational speed, present many problems for conventional cooling.
Various approaches have been proposed for utilizing internal fluid cooling to more effectively cool engine parts such as combustion chamber walls, turbine rotors and stators. In the case of rotor blades, some approaches have involved the internal use of a vaporizable cooling fluid that travels from the root of the rotor out through the tip of the rotor blade. Another approach has been to utilize a closed cycle cooling system in which a cooling fluid occupies only a portion of an internal cavity in the blade and circulates as a heat exchange medium. The physical properties of the cooling fluid are such that it is vaporized in certain regions of the cavity by virtue of the operating temperature prevailing in those regions during normal operation of the engine. U.S. Pat. No. 5,299,418 describes one particularly advantageous structure for closed circulation of a vaporizable liquid phase coolant within the cavity of a turbine blade. The improvement claimed in that patent involves a geometry for distributing coolant fairly uniformly over the inner surface of a blade in an axial-flow gas turbine so as to achieve a distributed cooling effect for the entire blade. Other constructions are claimed in U.S. Pat. Nos. 5,857,836 and 5,954,478.
While the ""418 patent illustrates an axial flow turbine with its characteristic blade shape, other forms of turbine have different configurations and pose different challenges to implementing effective coolant circulation. For example, there exists a need for better cooling systems in the relatively common turbine architecture utilizing a centrifugal compressor with an annular combustion chamber to feed a radial flow turbine. Similarly, smaller turbines where a regenerative loop architecture is used to enhance heat efficiency of a radial flow turbine, present particular challenges.
In addition to cooling of the turbine rotors that reside in combustion gases, cooling of compressor rotors is also desirable. Since the blades of the compressor operate under essentially adiabatic conditions in the air or gas mixture passing through the compressor, and the gas temperature rises as it is compressed, the ability of the blade material to operate at elevated temperature sets an upper limit on the temperature ratio, hence also the pressure ratio, of the compression system.
The compressor discharge temperature of gas turbines with adiabatic compression is typically limited to approximately 1200xc2x0 F. (920xc2x0 K) by the strength of available materials at temperature. For isentropic (and adiabatic) compression this implies a pressure ratio of 50 if the inlet temperature is 300xc2x0 K, and implies an ideal Brayton cycle efficiency of 0.67. Thus, there is considerable margin for improvement if the cycle pressure ratio, or the temperature ratio, can be raised. Moreover, operation of a compressor at elevated temperature lowers its efficiency, increasing the work required to achieve a given pressure ratio.
Accordingly, it is desirable to provide a system and construction for cooling the blades of a radial flow turbine so that the combustion process can be operated at higher temperatures without impairing the structural integrity of the turbine itself.
It is also desirable to provide a system and construction for cooling a rotor or blade assembly of a radial flow compressor to enhance its efficiency, permit use of less costly materials, or increase the pressure ratio for enhanced output.
In general, it is an object of the invention to provide an internal combustion engine wherein higher combustion temperatures or output power can be achieved while maintaining material temperatures at levels at least as low as those associated with known turbine engines and systems.
Another object of the invention is to provide an engine, having a radial flow centrifugal compressor, a combustion system and a radial flow turbine, that utilizes closed cycle evaporative cooling for the moving parts of the engine.
Still another object is to provide an improved rotor or rotor blade for use in a turbine or in a compressor of such a system.
One or more of the above desirable objects are achieved in accordance with the present invention by a system including, in a first aspect of the invention, a radial flow turbine having an arrangement of one or more stators and rotors in which each of the rotors defines an internal cavity that includes a vaporization section and a condensation section. The condensation section is disposed radially inward toward the shaft and the vaporization section extends over the rotor in thermal proximity to the blades. The vaporization section includes a series of pockets or passages for dispersing the cooling fluid proximate to heated surfaces of each blade, and a cascaded series of catchment channels or protruding shelves to distribute coolant to the pockets. A working turbine system includes a centrifugal compressor which feeds a compressed air/fuel mixture to an annular combustion chamber that, in turn, provides hot gases along a radial direction to impinge on the surface of a radial flow rotor. Optionally, the system is a regenerative system including a heat exchange sub-assembly which couples heat from the exhaust stream to a position between the compressor and combustion chamber.
In accordance with another aspect of the invention, a system includes a compression stage wherein an arrangement of one or more rotors compresses a fluid, such as air or a fuel mixture. The internal structure of the rotors includes a cavity with vaporization and condensation sections that operate to lower peak rotor temperature and transfer heat to a conveniently cooled sink. The vaporization section may include internally arranged shelves, lips or ledges that are arranged such that a coolant or phase change material cascades between successive catchment shelves, or is otherwise obstructed and channeled so as to distribute the coolant material effectively over the rotor interior and apply heat from the compressor blades to vaporize the coolant. This in turn drives a return flow or coolant circulation cycle in which heat is carried to a condensation region, lowering rotor operating temperature. The compressor may be a centrifugal compressor, such as a radial compressor that feeds a compressed fuel-air mixture to an annular combustion chamber for driving a radial-flow turbine, or it may be an axial flow compressor. In a turbine-driven power system, both the compressor and the turbine rotors may be cooled by internal evaporation, allowing enhancement of operating efficiency, power output and materials. Fully-cooled turbines may thus be manufactured at low cost, at any level of size or complexity, using low density high strength materials such as aluminum alloy that are otherwise inappropriate for conventional turbine or compressor operating temperatures.
In a further embodiment, the compressor, the turbine or both the compressor and turbine, of the present invention can be retrofitted between a compressor and a turbine of an existing axial-flow turbine power plant to enhance power or operating efficiency, essentially interfitting a high pressure, high temperature core engine in an otherwise conventional gas turbine system.